JPH094845A - Combustion apparatus - Google Patents

Abstract

PURPOSE: To realize the operation at high temperature rise level and the re- ignition at super-high altitude by providing a part of an opening of a dome and a part of fixed opening rows of the dimensions by which the reduced air level in the prescribed space in the combustion region is smaller than the level outside the space. CONSTITUTION: An annular combustion region 16 is demarcated by an outer looper liner 18 and an inner looper liner 20 to demarcate annular passages 22, 24. A part of the air flows through an opening 30, cools the liners, and a part of the air is sucked into a combustor. Other part of the air is passed through the passages 22, 24 to cool the combustion gas. A part of the air is passed through an opening 38 in a dome 40, and the opening can be of the shape of a vortex flow nozzle formed in a fuel nozzle. The space for low air flow in the combustor receives the air through an air opening 38 and a combustion air hole 32 of such dimensions as to reduce the air flow compared with the air hole which is not small in section.

Description

Description: FIELD OF THE INVENTION The present invention relates to gas turbine engines and in particular to annular burners.

(Prior Art) As is well known, combustors for gas turbine engines have made significant technological advances over the last few years from a technical point of view. As for combustion efficiency, for example, aircraft jet engines operate at high operating efficiencies in the 90% range. In addition, the technology has made significant improvements in reducing or eliminating the generation of pollutants and smoke from the combustion process.

 However, further advances in combustor technology are needed for combustors seeking to meet future aircraft demands. Clearly, it can be said that aircraft engine performance depends on obtaining higher turbine inlet temperatures. Higher turbine inlet temperatures within defined limits improve the estimated-to-weight ratio and the fuel consumption rate, which in turn leads to improved engine performance.

 Therefore, future demands require that new combustor arrangements for these high performance engines must operate at higher temperature rises at significantly higher power than modern combustors. However, it is of utmost importance for combustor designers to meet the requirements of high temperature rise without compromising the traditionally established levels of smoke emission from engines.

 In addition to the stringent requirements already mentioned, the combustor must be able to reignite within a defined altitude range. Also, when these engines are used in lightweight aircraft, they will operate below the level of temperature rise associated with modern combustors during engine deceleration and idling, i.e., idle, so they are sufficiently stable to facilitate ground maneuverability. It seems that something is necessary.

 It may be worth comparing the combustor requirements with the stability of the combustion process to gain a deeper insight into the complexity and operability issues associated with hot combustors. For this reason, FIG. 1 shows a plot of the relationship between the fuel / air ratio and the stability correlation parameter. As is known, most combustion occurs in the primary region or area of the combustor and this section of the combustor will be chosen for consideration. The standardized stability correlation parameters in combustor technology include the following terms applied to the primary region.

V = average flow-through velocity P = pressure level Y = inlet air temperature These terms are combined within the dimensional parameter (V / PT ² ) to increase velocity, reduce pressure or decrease temperature to increase stability parameter Increase to level. Similarly, they can adversely affect the combustion process, making it more difficult to continue combustion.

In other words, the higher value of the stability parameter indicates that the combustor operating requirements are more stringent and difficult.

 As discussed with reference to FIG. 1, this curve defines a stability limit that is produced by decreasing or increasing the fuel / air ratio until blowout occurs.

Therefore, the combustion condition in the range on the left side of the parabolic curve A is stable, and the combustion condition in the range on the right side of the curve A is unstable. Curve B defines the upper limit of the fuel / air ratio, and operation above this limit represents the extent of excessive smoke emission.

 Therefore, as is apparent from the above-described combustor operation for conventional modern combustors, the fuel / air ratio in the primary region is such that the fuel / air ratio falls within the range below curve B and the engine The deceleration, idling and high altitude re-ignition requirements of the vehicle are set to fall within the left side of curve A (stable operation). This is shown by curve C, where altitude reignition operation is defined as an operating point E with a reference altitude of about 9150 m (30,000 feet) and using the aircraft flight speed Mach number of 0.8 as the reignition point.

 Curve F shows the behavior of the engine at the higher temperature rise levels considered for future new technology engines. Clearly, to ensure combustion operation below the smoke limit (curve B), the combustor primary region airflow must also be increased so that the combustor operating curve F approaches the stability limit curve A as closely as possible. Notably, this clearly indicates that the operating requirements of the combustor have increased. Further, as described above, the deceleration and the high altitude reignition indicated by the point G can be set within the range outside the stability limit (curve A).

What is even more apparent is that this trend shifts the operating curve of the combustor towards increasingly stringent demands with increasing values of temperature rise, even with idle rotations (H and J for curves C and F respectively). Also outside the stability limit (curve A).

 Engineers and scientists have been fighting this problem for some time now and have tried to solve it in several different ways. These methods, all of which have substantial disadvantages, include variable fuel stage feed, variable geometry, and dual annular combustors. There is no known known solution to the above problem in the prior art.

 In the staged supply of fuel, considering that more fuel flows through some nozzles than other nozzles and fuel flows through two or more types of fuel nozzles, fuel is sent to the combustor through fuel nozzles. There is.

Therefore, if properly determined, the fuel is locally concentrated in a part of the combustor, so that the deceleration and idling conditions of the curve F are within the stability limit as shown by the dotted line K extending from the bottom of the curve F. Can be maintained. These schemes do not solve the basic stability requirement, and the problem of high reignition still remains, as indicated by the point G in the graph of FIG.

 A variable geometry combustor is a more realistic solution to the problems listed above, but at the cost of additional cost, weight and complexity associated with the hardware required to achieve variable geometry. Only possible. In this measure, the air flow to the primary region is mechanically altered by adjusting the air conditioning orifice. The fuel / air ratio and stability parameters at the combustor operating point determined accordingly are varied so that the operating demand remains in the stable region to the left of curve A.

 The last measure proposed is a double annular combustor, which uses a stepped air flow to solve the above problems. An example of a dual annular combustor is disclosed in U.S. Pat. No. 3,934,409 issued January 27, 1976 to H.A. Quillevere et al.

 In these combustors, the primary region consists of an inner annulus and an outer annulus. One of these annuli (first order) is designed to accept a relatively low air flow which exhibits good stability properties. The other annulus (secondary) accepts relatively high airflow and is allowed to exceed stability limits at low power and high reignition operating conditions. Assuming that the primary annulus sustains combustion and transfers it to the secondary annulus, combustion is sustained and the combustor is reignitable at high altitude conditions.

 However, the double annular structure inherently requires a separate air intake and fuel nozzle, resulting in a heavier and more expensive combustor.

 The measures of this kind mentioned above as well as the state-of-the-art combustion devices of the conventional device all give a uniform air flow distribution around. In some cases the air distribution is locally regulated in some way around the fuel nozzles or diffuser struts to give a uniform outlet temperature distribution. An example of locally controlling the air flow around a fuel nozzle is disclosed in U.S. Pat. No. 4,696,157 granted to G.Y.G. Barbier et al. On September 29, 1987. This patented device, as well as the other devices listed here, repeats the airflow distribution around the combustor, which is a function of the fuel nozzle or diffuser strut when considering a diffuser strut. Is.

 In all of these cases, the designer attempts to create an airflow distribution that is even or repeatable around the periphery. Also, during high-power operation, the fuel flow distribution is uniform for each nozzle.

 Another measure worth mentioning here was published in US Pat. No. 4,720,970 issued Jan. 26, 1988 to D.A. Hudson et al., And is similar to the variable geometry measures described above. In the above-referenced U.S. Pat. No. 4,720,970, the annular combustor dome is partitioned along its perimeter into one or more compartments. An airflow control valve alters the airflow entering a defined compartment (s) to define an area in the primary zone and varies the fuel / air ratio in that compartment (s) to control combustion . However, like the deformable measures described above, this measure also has the same disadvantages.

 We have found that inadequate extinguishment and stability limits of combustors can be extended, except for the disadvantages we have detailed above.

(Problem to be Solved by the Invention) An object of the present invention is to provide an improved annular combustor for a gas turbine engine capable of operating at a high temperature rise level and capable of high altitude reignition. .

(Means for Solving the Problems) One of the features of the present invention is that it has one section having a different air flow distribution and divides the entire engine into a peripheral area that maintains a stable operating range under operating conditions. To provide an improved combustor.

 Yet another feature of the present invention is that it allows existing combustor technology and hardware to be utilized while minimizing combustor system complexity, size, weight and cost. is there.

 The above and other features and advantages of the present invention will become more apparent from the following description and the accompanying drawings.

EXAMPLE In that example, the present invention is described with respect to a conventional single annular combustor, but for those skilled in this path it could equally well be used with any multi-nozzle combustor. Although the annular combustor is described here as a conventional annular burner liner made from sheet metal, the particular construction of the combustor and the details of that liner can be any other known construction as well.

 In FIG. 2, a conventional combustor, generally designated 10, having an outer combustor case 12 and an outer combustor case 14 and suitably coupled to other engine case members (not shown) is shown. It is shown. Details of the gas turbine engine are omitted here for simplicity and convenience of explanation, but the details are not fully assigned to the Platt of United Technologies Corp., the assignee of the present invention, all of which are incorporated herein by reference. Reference should be made to engine models that use annular combustors, such as those in the JT9D and F100 series of engines involved in the manufacture of the And Whitney section.

 Annular combustion region 16 is defined by an outer louver liner 18 and an inner louver liner 20, both of which are isolated from adjacent combustor case members 12 and 14, respectively, and together define annular passages 22 and 24. Cooled air from an engine compressor (not shown) flows through a portion of the air formed by the louver liner 30 to film cool the liner and another portion of the air. The pre-diffuser 26, dump diffuser, so that it is taken into the combustor via the combustion holes 32 for combustion purposes and a further part is taken in to cool the combustion gases (through the dilution air holes 34). A portion of the air passed through the passages 22 and 24 via 28 is passed through an opening 38 formed in the dome 40, which takes the form of a swirl nozzle built into the fuel nozzle. It can be done. The dimensions of all of these openings determine the distribution of axial and circumferential airflow entering the combustor and define the fuel / air ratio of the primary region formed adjacent dome 40.

 Fuel is passed to the combustor through any suitable fuel nozzle, generally indicated at 42. A plurality of fuel nozzles (only one is shown) 42 are arranged at equal intervals in the circumferential direction and are supported in the dome 40.

 In accordance with the present invention, at least one compartment in the primary region operates in stable conditions during the entire combustor envelope. Therefore, in order to withstand a higher temperature rise, the remaining area on the circumference has a higher air distribution than the other areas and the air flow around the circumference is uneven.

 Similarly, the fuel flow to the combustor will also be non-uniform on the circumference, balancing the distribution of the air flow and having a uniform pattern at the combustor exit. To achieve this, the size of the air orifice in the dome is adjusted to meet the desired air flow requirements. Similarly, the dimensions of the fuel nozzle orifices are also chosen to provide the required fuel distribution. From the above, a conventional annular combustor can be easily retrofitted to utilize the present invention.

 The flow distribution is best understood with reference to FIG. 3 which schematically shows the air openings on the outer circumference of the combustor.

The air opening 38 corresponds to the air opening 38 in FIG.

As mentioned above, the present invention considers the low airflow section in the combustor identified by the reference character s. The graph depicted in Figure 3 shows the distribution of air around the circumference of the annular combustor. Curve T shows the air flow distribution, the hanging part of the curve T shows the air flow in the low air flow compartment, which is dimensioned to reduce the air flow compared to the non-thinned air holes. Air is received through the portion of the air opening 38 and the portion of the combustion chamber pore 32 that are open. As is clear from the above, the flow distribution in the s section is reduced from the flow around the combustor. The fuel flow nozzles, aligned with these air openings, are also dimensioned to provide the fuel flow absolutely necessary to achieve the desired fuel / air ratio.

The present invention considers more than one compartment, for example, where it is desirable to place a low airflow compartment adjacent each firearm of the combustor to ensure stable operation at the location of ignition. There is. .

 The meaning of having compartments as shown by the present invention is best understood with reference to FIG. 4, which is the same graph as FIG. As mentioned above, the curve X represents the operation of the combustor in the low airflow compartment and the curve Y represents the operation of the combustor in the other compartments.

As is clear from the figure, the ignition point X'in the low air flow section enters the stable operation area, but the ignition point Y'in the high air flow section goes out of the stable operation area.

Since the ignition point in the low air flow section is within the stable operation range, it is possible to continue combustion and spread the combustion to the high air flow section.

 To achieve the above-mentioned uniformity, the fuel flow is proportioned or proportioned to the air flow in each compartment to provide a uniform outlet fuel / air ratio or temperature rise. This not only enhances engine operation but also extends turbine life. Obviously, low fuel flow may be achieved, for example, by reducing flow schedules in the region of low air flow, or by using a fuel controller to separately regulate the flow of fuel to the two regions. Can be obtained in several ways. However, this embodiment contemplates using differently sized fuel nozzles.

 The fuel flow schedule is shown in FIG. 5, where curve AA shows the fuel flow schedule in the high air flow region and curve BB shows the fuel flow in the low air flow region. It is also contemplated within the scope of the present invention that fuel flow scheduling is combined with the above-described fuel staging, especially when fuel flow is low in the low air flow region. ing. It is desirable to increase the fuel flow rate in this region of operation.

 As is apparent from the above, the present invention allows for use at high temperatures while avoiding the cost, weight and stability constraints mentioned in connection with the dual ring configuration.

 Although the present invention has been shown and described with respect to detailed examples thereof, those skilled in the art can make various changes in the shapes and details without departing from the spirit and scope of the claimed invention. It will be understood.

[Brief description of drawings]

 FIG. 1 shows the combustor operation, stability limit and smoke limit in the primary region of the combustor, plotting the fuel / air ratio against stability parameters, and FIG. 2 incorporates the present invention. Fig. 3 is a partial view and a schematic view showing a partial cross section of a conventional combustor that can take various forms. Fig. 3 shows the arrangement of fuel nozzles and the air flow around the dome of the combustor. Fig. 4 and Fig. 4 are the same as the graph of Fig. 1, but showing the combustor working liner of the present invention, and Fig. 5 is a graph showing the fuel flow in the fuel nozzle in the partitioned dome of the combustor. FIG. 6 is a diagram in which the fuel flow (Wf) is plotted against the pressure drop (ΔP) in the nozzle. 10: Annular combustor 12,14: Combustor case 16: Annular combustion area 18,20: Louver liner 22,24: Annular passage 30,38: Opening 40: Dome 42: Fuel nozzle

Claims (4)

[Claims] 1. An annular combustor for a gas turbine, comprising: an elongated outer annular liner and an elongated inner annular liner, and an annular combustor supported on the outer liner and the inner liner, all defining a combustion zone. A dome surrounding the front end, a plurality of openings spaced around the circumference of the dome for passing air through the combustion area, and air passing radially through the combustion area A portion of the plurality of openings in the dome and the fixed opening, the inner liner and at least one row of fixed openings in the outer liner being circumferentially spaced apart in a transverse plane. A portion of the rows of the above is sized so that the reduced air level in the defined compartment of the combustion zone is less than the level outside the compartment, under all operating conditions and in the compartment. An annular combustor portion outside the even When operating under the conditions are not stable, characterized in that the partition is operated stably. 2. A plurality of fuel nozzles mounted on the dome and spaced around the periphery, and the opening in the dome having an air swirler in the fuel nozzle. Annular combustor. 3. An annular combustor for a gas turbine, comprising: an outer annular liner and an inner annular liner that are concentrically located with respect to each other and define an annular combustion area; and a dome closing one end of the fuel area. Means attached to the inner and outer liners; a plurality of circumferentially spaced openings in the dome for passing air into the combustion zone; and the inner annular liner for passing air into the combustion zone. And at least one row of circumferentially spaced fixed openings provided in the outer annular liner and axially arranged with respect to the dome, and the gas turbine engine for supplying fuel into the combustion region. A plurality of fuel nozzles located within the dome for generating combustion products for use within the dome and a portion of the opening in the dome in a reduced area. The flow rate of the fuel and the defined section formed in the combustion region by providing a part of the fixed opening in the inner annular liner and the outer annular liner in the portion where the area is provided and the area is reduced. An annular combustion with a device for presetting the flow rate of the fuel to match the air flow distribution within the compartment and within the area outside the compartment to obtain an even fuel / air ratio over the combustion region. vessel. 4. A device for scheduling the flow rate of fuel such that the amount of fuel delivered to said compartment is increased to a high value during operation of the combustor which is being delivered with the least amount of fuel. The annular combustor according to claim 3, wherein